Typically, gas turbine engines include a compressor for compressing air, a combustor for mixing the compressed air with fuel and igniting the mixture, and a turbine blade assembly for producing power. Combustors often operate at high temperatures that may exceed 2,500 degrees Fahrenheit. Typical turbine combustor configurations expose turbine blade assemblies to these high temperatures. As a result, turbine blades must be made of materials capable of withstanding such high temperatures. In addition, turbine blades often contain cooling systems for prolonging the life of the blades and reducing the likelihood of failure as a result of excessive temperatures.
Typically, turbine blades, as shown in FIG. 1, are formed from a root portion at one end and an elongated portion forming a blade that extends outwardly from a platform coupled to the root portion at an opposite end of the turbine blade. The blade is ordinarily composed of a tip opposite the root section, a leading edge, and a trailing edge. The inner aspects of most turbine blades, as shown in FIG. 2, typically contain an intricate maze of cooling channels forming a cooling system. The cooling channels in the blades receive air from the compressor of the turbine engine and pass the air through the blade. The cooling channels often include multiple flow paths that are designed to maintain all aspects of the turbine blade at a relatively uniform temperature. However, centrifugal forces and air flow at boundary layers often prevent some areas of the turbine blade from being adequately cooled, which results in the formation of localized hot spots. Localized hot spots, depending on their location, can reduce the useful life of a turbine blade and can damage a turbine blade to an extent necessitating replacement of the blade.
Some conventional turbine blades incorporate serpentine cooling channels for directing cooling fluids through internal aspects of a turbine blade. Often times, the channels forming the cooling channels are nearly equal in cross-sectional area. The cooling channel proximate to the leading edge has a chordwise cross-section with a generally triangular shape. The apex of the triangular shaped cooling channel is the leading edge of the turbine blade. The configuration of the cross-sectional area negatively affects the distribution of cooling fluids to the leading edge and reduces the cooling fluid flow velocity as well as the internal heat transfer coefficient.
Other conventional cooling systems have attempted to overcome the negative impacts of the shape of the cross-section of the leading edge cooling channel by decreasing the size of the leading edge cooling channel relative to the downstream return cooling channel, as shown in FIG. 2. In short, the central rib has been shifted closer to the leading edge, thereby resulting in a leading edge cooling channel having a reduced cross-sectional area. The reduced cross-sectional area in the leading edge cooling channel increases the velocity of the cooling fluids, but causes the separation of cooling fluid flow in the tip region and a temperature increase at the blade tip. Therefore, while the reduced cross-sectional area of the leading edge cooling channel reduces the temperature at the leading edge, the temperature in the tip region has increased. Thus, a need exists for a cooling system for a turbine blade with a serpentine cooling channel that has increased heat transfer capabilities.